EP0659234B1 - Shaft power transfer in gas turbine engines - Google Patents
Shaft power transfer in gas turbine engines Download PDFInfo
- Publication number
- EP0659234B1 EP0659234B1 EP94919761A EP94919761A EP0659234B1 EP 0659234 B1 EP0659234 B1 EP 0659234B1 EP 94919761 A EP94919761 A EP 94919761A EP 94919761 A EP94919761 A EP 94919761A EP 0659234 B1 EP0659234 B1 EP 0659234B1
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- European Patent Office
- Prior art keywords
- engine
- gas turbine
- spool
- engine according
- spools
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
- F02C3/113—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission with variable power transmission between rotors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/32—Arrangement, mounting, or driving, of auxiliaries
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This invention relates to gas turbine engine transmission systems.
- the invention concerns transmission systems suitable for effecting power transfer between independently rotatable engine shafts in multispool gas turbine engines.
- a two-spool gas turbine engine arrangement in accordance with the pre-characterising of claim 1 appended hereto is known from GB Patent No 971,690.
- This engine arrangement comprises a high pressure spool consisting of a high pressure turbine section joined by a connecting shaft to a high pressure compressor section.
- a second, low pressure spool consisting of a low pressure turbine section and a low pressure compressor section are connected by a second shaft disposed concentrically with the first shaft.
- a third shaft is mounted concentrically with the other two shafts and is coupled to the low pressure spool shaft in order to act alternatively as a starter motor power input or as an auxiliary generator power offtake.
- the high and low pressure spools are coupled one to the other only by means of the engine gas path.
- the starter motor/auxiliary generator unit comprises a power transfer machine coupled to the low pressure engine spool and is capable of operating in forward or reverse power transfer mode according to whether the unit is to be operated as a starter motor or as a generator.
- this power transfer means is coupled to the single spool only and to an external connection, and contains no provision for power transfer between the spools of the engine.
- a multispool gas turbine 10 having a conventional transmission system is shown schematically in Figure 1.
- the gas turbine engine shown comprises, in flow series, a front fan assembly 12 and a core engine 14.
- the engine is of the ducted fan by-pass type and has three relatively rotatable spools including a low pressure spool 16, an intermediate pressure spool 18, and a high pressure spool 20.
- the low pressure spool includes a fan 12, a multistage turbine assembly 22 located at the downstream end of the core engine, and an interconnecting load transmitting shaft 24 rotatable about engine axis 26.
- the intermediate pressure spool 18 includes a multistage axial flow compressor 28, a turbine rotor assembly 30, and a hollow interconnecting shaft 32 concentrically disposed around engine shaft 24.
- the engine's high pressure spool 20 similarly includes a multistage axial flow compressor 34, a turbine rotor assembly 36, and an interconnecting shaft 38 concentric with engine shafts 32 and 24.
- the transmission includes a radial power off-take shaft 40, a so called step-aside gearbox 42 drivingly connected to the engine's high pressure spool by the drive shaft 40, an externally mounted accessory gearbox 44, and a drive shaft 46 connecting the accessory gearbox to the step-aside gearbox 42.
- Various accessories (not shown), both engine and aircraft, are mounted on the accessory gearbox 44 so as to be driven by the transmission during engine operation.
- This configuration is found in many ducted fan multispool gas turbine engines. It has the advantage over other configurations in that it allows the same transmission system to be utilised for transferring engine starter torque to the engine's high pressure spool during ground starting, as well as engine power to the accessories during self sustained operation. There is a drawback, however, with this type of arrangement.
- aircraft mounted gas turbine engines typically rely upon the free rotation of the engine spools to generate sufficient core engine flow to support combustion and rapid engine acceleration at re-light.
- this capability can be significantly reduced.
- the additional load imposed by the accessories reduces the free rotational speed of the spool, and as a result the airflow through the core engine.
- Another method involves transferring power to the high pressure spool of an extinguished engine from a source external to the engine. Generally this is achieved using bleed air from a neighbouring engine to drive the starter of the extinguished engine.
- the power transferred augments that of the freely rotating high pressure spool and as such causes the spool to rotate faster, improving the chances of successful re-light.
- a major drawback with this is that it relies on the continued functioning of at least one other engine.
- the present invention has for one of its objectives improvements to the in-flight re-light performance of aircraft mounted gas turbine engines, in particular the provision of a transmission system which is capable of transferring power from the engine's low pressure spool to it's high pressure spool following a combustion flame-out condition.
- Another object of the present invention is to improve engine part load performance.
- variable geometry is another way of achieving flow control at off-design conditions.
- variable angle aerofoils are provided in the compression systems of gas turbine engines. By rescheduling the aerofoil angles, the flow characteristics of individual compressor stages can be substantially altered.
- variable geometry flow control has a number of disadvantages. The additional hardware tends to add appreciably to the overall cost, weight and complexity of the engine.
- the present invention has for a second objective improvements to the part speed performance of multispool gas turbine engines, in particular the provision of a transmission system which is capable of re-scheduling the rotational speeds of individual engine spools at part speed operation.
- a further object of the present invention is to improve engine ground start performance.
- the invention has for a third objective improvements to the ground starting abilities of multispool gas turbine engines, in particular the provision of a transmission system which is capable of transferring engine starter torque to more than one of the engine spools during ground starting.
- the invention provides a gas turbine engine of the kind comprising at least two independently rotatable engine spools and a plurality of power transfer machines coupled to the engine spools including a first power transfer machine coupled to a first engine spool, said machine being capable of operation as a generator (forward power transfer) to take power from the spool or as a motor (reverse power transfer) to drive the spool, characterised in that there is provided at least on further power transfer machine coupled to another of the engine spools, which further machine is also capable of operation either as a generator (forward power transfer) to take power from the spool or as a motor (reverse power transfer) to drive the spool, power transmission means arranged to interconnected said plurality of power transfer machines and means to control the operation of said machines to transfer power selectively from at least one of the engine spools to at least one other of the engine spools.
- the gas turbine engine includes a plurality of engine driven accessories and the transmission means is adapted to transfer power from at least one of the engine spools to drive the accessories.
- the gas turbine engine is for mounting on an aircraft having a plurality of engine driven accessories and the transmission means is further adapted to transmit power from at least one of the engine spools to drive the aircraft accessories.
- a ducted fan gas turbine engine 10 is suspended from the wing 50 of an aircraft by a pylon 52.
- the engine 10 is of identical construction to that shown in Figure 1 having a front fan 12, which forms part of the low pressure spool 16 (Figure 1) and a core engine or gas generator 14, which contains the remainderof spool 16 and the two further independently rotatable spools 18 and 20 shown in Figure 1.
- Figure 2 shows the engine in part cut-away view, the majority of the engine detail having been omitted for clarity.
- the engine shown includes a transmission system 60 in accordance with a first general embodiment of the invention.
- the transmission comprises three hydraulic flow displacement machines 62, 64 and 66, each of which is drivingly connected to a respective one of the engine spools. All three machines are arranged in fluid supply communication to define a hydrostatic transmission system, the exact configuration of which will be described later.
- the machines 62, 64 and 66 are of the type which may be operated in a positive displacement mode as pump, converting a mechanical work input into a pressurised hydraulic output, and in a negative reverse mode as a motor, converting a pressurised hydraulic input into a mechanical work output. These machines are commonly found in hydrostatic transmission systems. Reversible flow swash plate type machines having a variable flow capacity are preferred in this embodiment for greater operational flexibility.
- a first of the flow machines 62 is mounted on the engine's step-aside gearbox 42. Together with radial drive shaft 40 the gearbox 42 connects the flow machine 62 to the engine's high pressure spool 20.
- the gearbox 42 is provided with bevel gearing (not shown) to turn the drive from shaft 40 through 90 degrees so that the flow machine 62, and its associated pipework (also not shown), can be located within the region 68 defined between core engine casing structure 70 and cowling 72.
- a second of the hydraulic flow machines 64 is drivingly connected to the engine's intermediate pressure spool 18.
- a second radial drive shaft 74 and a reduction gearbox 76 are provided to connect hydraulic machine 64 with the engine's intermediate pressure spool 18.
- gearbox 76 is mounted to core engine casing structure 70, together with hydraulic machine 64, inboard of cowling structure 72.
- gearbox 76 also includes bevel gearing (not shown) for turning the drive from shaft 74 through 90 degrees.
- a third of the hydraulic flow machines 66 is drivingly connected to the downstream end of the low pressure spool engine shaft 24.
- a reduction gearbox 78 is provided between the engine shaft 24 and the hydraulic machine 66.
- the gearbox is mounted to fixed engine structure 72 via supports 80 and vanes 82 positioned at the downstream end of the core engine. This arrangement is preferred since it provides for both aerodynamic positioning of the flow displacement machine 66 and operational accessibility.
- FIG. 3 shows schematically the layout of the hydrostatic transmission 60.
- the transmission shown may be integral with the engine oil system or separate if desire.
- each of the hydraulic machines 62, 64 and 66 are connected to a flow control 86 by separate hydraulic lines 88, 90 and 92 respectively.
- the flow control 86 has a first inlet 94 adapted to receive flow from machine 64 along line 90, a second inlet 96 adapted to receive flow from machine 66 along line 92, and an outlet 98 for delivering flow to machine 62 along line 88.
- machine 62 includes a first inlet 100 connected to a tank 102 by means of a hydraulic line 104 having a non-return valve 106 positioned within it, a second inlet 108 for receiving line 88, a first outlet 110 connected to machine 64 via a hydraulic line 112, a second outlet 114 connected to machine 66 via a hydraulic line 116, and a further outlet 118 connected to a drain line 120 to tank 102.
- the second machine 64 includes a first inlet 122 connected to tank 102 along a line 124 having a non return valve 125, a second inlet for receiving line 112, a first outlet connected to flow control 86 via line 90, and a second outlet 130 to the tank 102 via a drain line 132.
- the third machine 66 includes an inlet 134 connected to line 116 from machine 62 and an outlet 136 connected to the flow control 86 by line 92.
- the transmission system 60 is further provided with a cooler for cooling the working fluid passing along lines 92 and 112 during operation, and in addition includes shut-off valves 144 and 146 in lines 92 and 116 respectively for isolating machine 66.
- the transmission 60 may be configured to transfer power between selective engine spools.
- the transmission in a first mode the transmission is configured for power transfer between the low pressure and high pressure engine spool, the configuration that would be selected following an in-flight combustion flame-out condition.
- the hydraulic machine 66 connected to the engine's low pressure spool is configured to work as a pump, and the hydraulic machine 62 connected to the engine's high pressure spool as a motor powered by working fluid energised by the pump 66.
- the shut-off valves 144 and 146 open working fluid will pass from the pump 66 through the flow control 86 to power the motor 62.
- the flow control will ensure that hydraulic machine 64 is isolated from pressurised working fluid so that all available power will pass to the engine's high pressure spool.
- the transmission is configured for power transfer between the engine's intermediate pressure and high pressure spool, the configuration that would be selected for part speed operation.
- the hydraulic machine 64 connected to the engine's intermediate pressure spool is configured to work as a pump, and the hydraulic machine 62 connected to the engine's high pressure spool as a motor powered by working fluid energised by the pump 64.
- the flow control 86 isolates the hydraulic machine 66 so that all the pressurised working fluid passes to the motor 62 along lines 88 and 90 with the return flow passing back to the pump 64 along 112.
- the transmission is configured for power transfer between the engine's high pressure and intermediate pressure spool, the configuration that would be selected during ground starting.
- the hydraulic machine 62 connected to the engine's high pressure spool is configured to work as a pump, and the hydraulic machine 64 connected to the engine's intermediate pressure spool as a motor powered by working fluid energised by the pump 62.
- the flow control 86 isolates the hydraulic machine 66 so that all the pressurised working fluid passes to the motor 64 along line 112 with the return flow passing back to the pump 62 along lines 88 and 90.
- a further hydraulic flow displacement machine 148 is mounted in the pylon structure 52 and drivingly connected to a hydraulic pump 160 adapted to power the aircraft hydraulic systems. Collectively the flow displacement machine 148 and pump 160 define an engine to aircraft power transfer means.
- the flow displacement machine 66 is provided with an additional inlet 138 and an outlet 140 for connection to machine 148.
- a first hydraulic line 152 connects the outlet 140 to machine 148, and a second hydraulic line 150 connects machine 148 to the inlet 138 of machine 66 via a cooler 156.
- a third hydrostatic line 154 is provided to connect lines 150 and 152 to a tank.
- hydraulic machine 66 is permanently configured as a pump and machine 148 as a motor. At all times during engine operation the pump 66 delivers pressurised working fluid to the motor 148 to drive the aircraft hydraulic systems.
- transmission 60a to the engine of Figure 2 enables the aircraft hydraulic pump 160 to be positioned closer to the hydraulic systems it drives. By moving the pump 160 from the accessory gearbox 44 to the pylon structure 52 considerable pipework and hence weight can be removed from the engine.
- FIG. 4 shows the same gas turbine engine as Figure 2, but with a transmission system 60 in accordance with a second embodiment of the invention.
- the transmission shown connects only two of the engine spools, the high pressure spool and pressure spool 20 and 16 respectively.
- the transmission comprises first and second reduction gearboxes 162 and 164 drivingly connected to the engine's high and low pressure spools.
- Reduction gearbox 162 is connected to the engine's high pressure spool through step-aside gearbox 42, and reduction gearbox 164 to the engine's low pressure spool at the downstream end of shaft 24 as in the previous embodiment.
- Gearbox 164 includes bevel gearing (not shown) to transfer the drive from shaft 24 to a radial output shaft 166.
- Output shaft 166 is in two parts and extends from the gearbox 164 to a pylon mounted gearbox 168 through a differential 170.
- the axial differential drive 172 is of the type commonly used in aircraft electrical power generator systems. These drives are configured so that the majority of power passing through them is passed through a differential gear arrangement, whilst a small proportion is used to drive an integral variable speed hydrostatic transmission.
- the design of the differential drive is such that it allows two different speed inputs to be mechanically summed to produce a single speed output.
- the axial differential drive is arranged so that one of speed inputs is the mechanical input to the differential gear itself, and the other the output from the hydrostatic transmission.
- the drive ratio of the differential drive is infinitely variable by virtue of the variable hydrostatic transmission output.
- the transmission 60 of the second preferred embodiment operates as follows.
- the ratio of drive 172 essentially follows the ratio of low pressure spool to high pressure spool speed. The only time the drive 172 deviates from this is following an in-flight combustion flame out condition.
- the ratio of the drive 172 will alter in accordance with power transfer requirements. As previously discussed this will enable the engine's low pressure spool to accelerate the high pressure spool to the re-light condition.
- the transmission system 60 could be modified further so that power could be transferred selectively between any two of the engine spools. This could be achieved for example by adding a further axial differential drive gear between the engines high pressure 20 and intermediate pressure spool 18, or alternatively a hydrostatic transmission as in the first preferred embodiment.
- Figure 5 schematically represents the gas turbine engine of Figure 1 having a transmission system in accordance with a third embodiment of the invention.
- the high pressure spool 20 is adapted to drive the engine accessory gearbox 44 via drive shafts 40 and 46 and step aside gearbox 42 as before, and the radial drive shaft 40 is bevelled to the engine's high pressure spool at 180 in the usual way.
- the transmission shown in Figure 5 comprises a reduction gearbox 182 located in region 68 and drivingly connected to the engine's intermediate pressure spool 18, reduction and bevel gearing 184 drivingly connected to the engine's low pressure spool 16 at the downstream end of shaft 24, and bevel gearing 186 positioned radially outwards of gearing 184 in region 68.
- Gearbox 182 is connected to the engine's intermediate pressure spool 18 by a radial drive shaft 188 bevelled to shaft 32 at 190, and to step aside gearbox 42 by an axial drive shaft 192.
- gearing 186 is connected to gearing 184 by a radial drive shaft 194, and to step-aside gearbox 42 by an inclined drive shaft 196.
- a clutch assembly 198 is provided in drive shaft 192 so that the drive between the engine's high pressure spool 20 and intermediate pressure spool 18 may be selectively engaged and disengaged.
- a second clutch assembly 200 is provided in drive shaft 196 so that the drive between the engine's high pressure spool 20 and low pressure spool 16 may be selectively engaged and disengaged.
- engine re-light performance may be improved by engagement of clutch assembly 200 and disengagement of clutch assembly 198 following a flame-out condition.
- engine part speed performance may be improved by engagement of clutch assembly 198 and disengagement of clutch assembly 200 during part load operation, and ground starting performance by similar engagement and disengagement prior to engine ignition.
- a power transmission system 60 in accordance with a first aspect of a fourth embodiment of the invention is shown in Figure 6.
- Figure 6 shows schematically the gas turbine engine of Figure 1 having all the components of the prior art transmission.
- the transmission system shown also includes a number of additional components also found in the third embodiment of the invention.
- the transmission system of Figure 6 further comprises a first electrical machine 202 drivingly connected to the engine's accessory gearbox 44, a second electrical machine 204 drivingly connected to gearbox 182 and a third electrical machine 206 drivingly connected to the engine's low pressure spool via gearing 208.
- the electrical machines 202, 204 and 206 are of the type which may be operated in a forward mode as a generator, converting a mechanical work input into an electrical output, or in a reverse mode as a motor, converting an electrical input into a mechanical work output.
- the machines may be either permanent magnet or electromagnetic induction type machines.
- control 210 comprises switchgear adapted to isolate each of the machines 202, 204 and 206 from the transmission.
- power may be selectively transferred from the engine's low pressure spool 16 to it's high pressure spool 20 by operating machine 206 as a generator and machine 202 as a motor powered by current from the generator 206.
- power may be transferred from the engine's intermediate pressure spool 18 to it's high pressure spool 20 by operating machine 204 as a generator and machine 202 as a motor powered by the generator 204.
- power may be transferred from the engine's high pressure spool 20 to it's intermediate pressure spool 18 by operating machines 202 and 204 in reverse.
- FIG. 7 A second aspect of this embodiment of the invention is shown in Figure 7.
- the electrical induction machines 202, 204 and 206 are replaced together with the mechanical drive transmission, with switched reluctance electrical machines 222, 224 and 226 respectively.
- the switched reluctance machines are adapted top operate in both forward and reverse motor and generator modes.
- the advantage of the switched reluctance machines over the induction machines of the transmission of Figure 6 resides in their power to size ratio. In the arrangement shown this allows machines 222, 224 and 226 to be embedded within the engine internal of the gas flow parts.
- the rotors 228 of machines 222, 224 and 226 are integral with engine shafts 38, 32 and 24 respectively.
- the machines 222, 224 and 226 are connected electrically to control 210 along separate lines 212, 214 and 216 as before. Accordingly power may be transferred selectively between engine spools in the same way as in the embodiment of Figure 6.
- Figure 8 shows a power transmission system according to a fifth embodiment of the invention.
- Figure 8 shows schematically the gas turbine engine of Figure 1 having all the components of the prior art transmission.
- the transmission system shown also comprises an auxiliary air turbine 230 drivingly connected to the engine accessory gearbox 44.
- the auxiliary turbine 230 includes an inlet 232 which is in fluid flow communication with the downstream end of compressor 28.
- Ducting 234 links the inlet 232 of the turbine to an engine bleed flow means 236 positioned at the outlet to compressor 28.
- a bleed flow control valve 238 is positioned adjacent compressor bleed 236 so that compressor discharge air may be bled selectively from the engine to drive the turbine 232.
- bleed flow control valve 238 remains closed and the engine functions as normal. At part speed operation, however, bleed flow control valve 238 opens and compressor discharge air is fed to the turbine 230.
- the increase in effective discharge area causes the load on the turbine section 30 to increase the thereby the rotational speed of the intermediate pressure spool 18 to fall. This contrasts with an increase in the high pressure spool rotational speed due to the work input from turbine 230 via the engine accessory drive.
- the transmission system shown in this embodiment enables power to be selectively transferred from the intermediate pressure spool to high pressure spool by selective operation of valve 238.
- valve 238 the same effect could be achieved by adapting the turbine 232 to be driven by bleed gases bled from the inlet to turbine 30 instead of the outlet to compressor 28.
- opening the bleed to turbine 232 would effectively reduce the area ratio of the turbine 30 and thereby reduce the rotational speed of the spool 18 whilst at the same time increasing the rotational speed of spool 20.
- the arrangement shown in Figure 5 offers a number of advantages over such alternative arrangements.
- it enables the bleed flow generated as a result of flow control at part speed operation to be utilised in a more useful manner.
- the bleed flow is normally dumped to the bypass flow and as such has only a minor effect on cycle efficiency.
- the arrangement described however enables the bleed flow to be utilised to increase the flow capacity of the engines high pressure compressor 34 and thereby improve part speed engine performance.
Abstract
Description
- This invention relates to gas turbine engine transmission systems. In particular the invention concerns transmission systems suitable for effecting power transfer between independently rotatable engine shafts in multispool gas turbine engines.
- A two-spool gas turbine engine arrangement in accordance with the pre-characterising of claim 1 appended hereto is known from GB Patent No 971,690. This engine arrangement comprises a high pressure spool consisting of a high pressure turbine section joined by a connecting shaft to a high pressure compressor section. A second, low pressure spool consisting of a low pressure turbine section and a low pressure compressor section are connected by a second shaft disposed concentrically with the first shaft. A third shaft is mounted concentrically with the other two shafts and is coupled to the low pressure spool shaft in order to act alternatively as a starter motor power input or as an auxiliary generator power offtake. However, the high and low pressure spools are coupled one to the other only by means of the engine gas path. Also the starter motor/auxiliary generator unit comprises a power transfer machine coupled to the low pressure engine spool and is capable of operating in forward or reverse power transfer mode according to whether the unit is to be operated as a starter motor or as a generator. However, this power transfer means is coupled to the single spool only and to an external connection, and contains no provision for power transfer between the spools of the engine.
- A
multispool gas turbine 10 having a conventional transmission system is shown schematically in Figure 1.
The gas turbine engine shown comprises, in flow series, afront fan assembly 12 and acore engine 14. The engine is of the ducted fan by-pass type and has three relatively rotatable spools including a low pressure spool 16, an intermediate pressure spool 18, and a high pressure spool 20. The low pressure spool includes afan 12, amultistage turbine assembly 22 located at the downstream end of the core engine, and an interconnectingload transmitting shaft 24 rotatable aboutengine axis 26. The intermediate pressure spool 18 includes a multistageaxial flow compressor 28, aturbine rotor assembly 30, and ahollow interconnecting shaft 32 concentrically disposed aroundengine shaft 24. The engine's high pressure spool 20 similarly includes a multistageaxial flow compressor 34, aturbine rotor assembly 36, and an interconnectingshaft 38 concentric withengine shafts - The transmission includes a radial power off-take
shaft 40, a so called step-aside gearbox 42 drivingly connected to the engine's high pressure spool by thedrive shaft 40, an externally mountedaccessory gearbox 44, and adrive shaft 46 connecting the accessory gearbox to the step-aside gearbox 42. Various accessories (not shown), both engine and aircraft, are mounted on theaccessory gearbox 44 so as to be driven by the transmission during engine operation. - This configuration is found in many ducted fan multispool gas turbine engines. It has the advantage over other configurations in that it allows the same transmission system to be utilised for transferring engine starter torque to the engine's high pressure spool during ground starting, as well as engine power to the accessories during self sustained operation. There is a drawback, however, with this type of arrangement.
- Following an in-flight combustion flame out condition, aircraft mounted gas turbine engines typically rely upon the free rotation of the engine spools to generate sufficient core engine flow to support combustion and rapid engine acceleration at re-light. In arrangements where the accessories are driven by the engine's high pressure spool this capability can be significantly reduced. During periods of extinguished engine operation the additional load imposed by the accessories reduces the free rotational speed of the spool, and as a result the airflow through the core engine.
- One way of improving the re-light characteristics of ducted fan gas turbine engines is disclosed in our co-pending International Patent Application GB92/01179. This patent application discloses an aircraft mounted gas turbine engine having an accessory drive linked to the engine's low pressure spool. The arrangement ensures that in the event of a flame-out condition the engine's windmilling fan continues to drive the accessory gearbox. Under these conditions the power available from the windmilling is significantly greater than that required by the accessories and as such the accessory load impacts far less on the core engine flow.
- Another method involves transferring power to the high pressure spool of an extinguished engine from a source external to the engine. Generally this is achieved using bleed air from a neighbouring engine to drive the starter of the extinguished engine. The power transferred augments that of the freely rotating high pressure spool and as such causes the spool to rotate faster, improving the chances of successful re-light. A major drawback with this is that it relies on the continued functioning of at least one other engine.
- One way of overcoming this would be to utilise power from the windmilling fan of the extinguished engine, instead of bleed flow energy from the neighbouring engine, to accelerate the engine spool.
- Accordingly the present invention has for one of its objectives improvements to the in-flight re-light performance of aircraft mounted gas turbine engines, in particular the provision of a transmission system which is capable of transferring power from the engine's low pressure spool to it's high pressure spool following a combustion flame-out condition.
- Another object of the present invention is to improve engine part load performance.
- Since the aerodynamic characteristics of gas turbine engine compressor systems are designed for optimum operation at a particular condition, it is often necessary to provide some form of airflow control to avoid flow breakdown at part speed operation. This is usually achieved using compressor bleed wherein a portion of compressor flow is bleed from the engine to avoid compressor surge during part speed operation. Although effective in controlling the delivery flow pressure of individual compressor stages, bleed flow control has a number of drawbacks, the main disadvantage being its impact on cycle efficiency. Variable geometry is another way of achieving flow control at off-design conditions. Typically variable angle aerofoils are provided in the compression systems of gas turbine engines. By rescheduling the aerofoil angles, the flow characteristics of individual compressor stages can be substantially altered. Again although effective, variable geometry flow control has a number of disadvantages. The additional hardware tends to add appreciably to the overall cost, weight and complexity of the engine.
- It is well known that similar control could be achieved by re-scheduling the rotational speed of the individual engine spools at part speed operation. For example, by respectively increasing and decreasing the rotational speeds of the high and intermediate pressure spools of a three spool engine, cycle efficiency at part speed operation could be substantially improved. Unfortunately this is not possible with conventional engines since the rotational speed of each engine spool is determined by the fixed aerodynamics of the turbine driving it.
- Accordingly the present invention has for a second objective improvements to the part speed performance of multispool gas turbine engines, in particular the provision of a transmission system which is capable of re-scheduling the rotational speeds of individual engine spools at part speed operation.
- A further object of the present invention is to improve engine ground start performance.
- Conventionally ground starting of multispool gas turbine engines is effected by transmitting torque to the engine's high pressure spool via the engine accessory gearbox. Once sufficient core flow is developed ignition will take place. In general the time to ignition will depend on factors such as accessory loading and engine aerodynamics. In future engine applications there is a possibility that this period could increase, when compared with todays standards, as more and more demands are made on the accessory drive systems.
- Accordingly, the invention has for a third objective improvements to the ground starting abilities of multispool gas turbine engines, in particular the provision of a transmission system which is capable of transferring engine starter torque to more than one of the engine spools during ground starting.
- In its broadest sense the invention provides a gas turbine engine of the kind comprising at least two independently rotatable engine spools and a plurality of power transfer machines coupled to the engine spools including a first power transfer machine coupled to a first engine spool, said machine being capable of operation as a generator (forward power transfer) to take power from the spool or as a motor (reverse power transfer) to drive the spool, characterised in that there is provided at least on further power transfer machine coupled to another of the engine spools, which further machine is also capable of operation either as a generator (forward power transfer) to take power from the spool or as a motor (reverse power transfer) to drive the spool, power transmission means arranged to interconnected said plurality of power transfer machines and means to control the operation of said machines to transfer power selectively from at least one of the engine spools to at least one other of the engine spools.
- Preferably the gas turbine engine includes a plurality of engine driven accessories and the transmission means is adapted to transfer power from at least one of the engine spools to drive the accessories.
- Preferably the gas turbine engine is for mounting on an aircraft having a plurality of engine driven accessories and the transmission means is further adapted to transmit power from at least one of the engine spools to drive the aircraft accessories.
- The invention will now be described in greater detail with reference, by way of example only, to the accompanying drawings, in which:
- Figure 1 is a schematic view of a multispool gas turbine engine incorporating a transmission system of the prior art,
- Figure 2 is a partially sectioned partially cut-away side view of a gas turbine engine similar to that of Figure 1, but incorporating a transmission system according to a first embodiment of the invention,
- Figure 3 schematically represents the transmission system of the first embodiment of the invention,
- Figure 4 shows the same multispool gas turbine engine as Figure 2 incorporating a transmission system according to a second embodiment of the invention,
- Figure 5 schematically represents a multispool gas turbine engine similar to that of Figure 2 incorporating a transmission system according to a third embodiment of the invention,
- Figure 6 shows the same multispool gas turbine engine as Figure 5 incorporating a transmission system according to one aspect of a fourth embodiment of the invention,
- Figure 7 shows the same multispool gas turbine engine as Figure 5 incorporating a transmission system according to another aspect of the fourth embodiment of the invention, and,
- Figure 8 shows the same multispool gas turbine engine as Figure 5 incorporating a transmission system according to one aspect of a fifth embodiment of the invention, similar parts having the same reference numerals throughout.
- Referring to Figure 2, a ducted fan
gas turbine engine 10 is suspended from thewing 50 of an aircraft by apylon 52. - The
engine 10 is of identical construction to that shown in Figure 1 having afront fan 12, which forms part of the low pressure spool 16 (Figure 1) and a core engine orgas generator 14, which contains the remainderof spool 16 and the two further independently rotatable spools 18 and 20 shown in Figure 1. Unlike Figure 1, Figure 2 shows the engine in part cut-away view, the majority of the engine detail having been omitted for clarity. - The engine shown includes a
transmission system 60 in accordance with a first general embodiment of the invention. Essentially the transmission comprises three hydraulicflow displacement machines - The
machines - With reference to Figure 2, a first of the
flow machines 62 is mounted on the engine's step-aside gearbox 42. Together withradial drive shaft 40 thegearbox 42 connects theflow machine 62 to the engine's high pressure spool 20. Thegearbox 42 is provided with bevel gearing (not shown) to turn the drive fromshaft 40 through 90 degrees so that theflow machine 62, and its associated pipework (also not shown), can be located within theregion 68 defined between coreengine casing structure 70 andcowling 72. In a similar manner a second of thehydraulic flow machines 64 is drivingly connected to the engine's intermediate pressure spool 18. A secondradial drive shaft 74 and areduction gearbox 76 are provided to connecthydraulic machine 64 with the engine's intermediate pressure spool 18. Like step-aside gearbox 42gearbox 76 is mounted to coreengine casing structure 70, together withhydraulic machine 64, inboard ofcowling structure 72. In addition to reduction gearing,gearbox 76 also includes bevel gearing (not shown) for turning the drive fromshaft 74 through 90 degrees. A third of thehydraulic flow machines 66 is drivingly connected to the downstream end of the low pressurespool engine shaft 24. Areduction gearbox 78 is provided between theengine shaft 24 and thehydraulic machine 66. The gearbox is mounted to fixedengine structure 72 viasupports 80 andvanes 82 positioned at the downstream end of the core engine. This arrangement is preferred since it provides for both aerodynamic positioning of theflow displacement machine 66 and operational accessibility. - Referring now to Figure 3 which shows schematically the layout of the
hydrostatic transmission 60. The transmission shown may be integral with the engine oil system or separate if desire. As shown each of thehydraulic machines flow control 86 by separatehydraulic lines flow control 86 has afirst inlet 94 adapted to receive flow frommachine 64 alongline 90, asecond inlet 96 adapted to receive flow frommachine 66 alongline 92, and anoutlet 98 for delivering flow tomachine 62 alongline 88. In alike manner machine 62 includes afirst inlet 100 connected to atank 102 by means of ahydraulic line 104 having anon-return valve 106 positioned within it, asecond inlet 108 for receivingline 88, afirst outlet 110 connected tomachine 64 via ahydraulic line 112, asecond outlet 114 connected tomachine 66 via ahydraulic line 116, and afurther outlet 118 connected to adrain line 120 totank 102. Thesecond machine 64 includes afirst inlet 122 connected totank 102 along aline 124 having anon return valve 125, a second inlet for receivingline 112, a first outlet connected to flowcontrol 86 vialine 90, and asecond outlet 130 to thetank 102 via adrain line 132. Thethird machine 66 includes aninlet 134 connected to line 116 frommachine 62 and anoutlet 136 connected to theflow control 86 byline 92. Thetransmission system 60 is further provided with a cooler for cooling the working fluid passing alonglines valves lines machine 66. - During engine operation the
transmission 60 may be configured to transfer power between selective engine spools. With reference to Figure 3, in a first mode the transmission is configured for power transfer between the low pressure and high pressure engine spool, the configuration that would be selected following an in-flight combustion flame-out condition. In this mode of operation thehydraulic machine 66 connected to the engine's low pressure spool is configured to work as a pump, and thehydraulic machine 62 connected to the engine's high pressure spool as a motor powered by working fluid energised by thepump 66. When the shut-offvalves pump 66 through theflow control 86 to power themotor 62. In this mode the flow control will ensure thathydraulic machine 64 is isolated from pressurised working fluid so that all available power will pass to the engine's high pressure spool. - In a second mode the transmission is configured for power transfer between the engine's intermediate pressure and high pressure spool, the configuration that would be selected for part speed operation. In this mode the
hydraulic machine 64 connected to the engine's intermediate pressure spool is configured to work as a pump, and thehydraulic machine 62 connected to the engine's high pressure spool as a motor powered by working fluid energised by thepump 64. In this mode theflow control 86 isolates thehydraulic machine 66 so that all the pressurised working fluid passes to themotor 62 alonglines pump 64 along 112. - In a third mode the transmission is configured for power transfer between the engine's high pressure and intermediate pressure spool, the configuration that would be selected during ground starting. In this mode the
hydraulic machine 62 connected to the engine's high pressure spool is configured to work as a pump, and thehydraulic machine 64 connected to the engine's intermediate pressure spool as a motor powered by working fluid energised by thepump 62. In this mode theflow control 86 isolates thehydraulic machine 66 so that all the pressurised working fluid passes to themotor 64 alongline 112 with the return flow passing back to thepump 62 alonglines - In order to reduce engine weight the engine of Figure 2 is provided with an additional
hydrostatic transmission 60a. A further hydraulicflow displacement machine 148 is mounted in thepylon structure 52 and drivingly connected to ahydraulic pump 160 adapted to power the aircraft hydraulic systems. Collectively theflow displacement machine 148 and pump 160 define an engine to aircraft power transfer means. With reference now to Figure 3, theflow displacement machine 66 is provided with anadditional inlet 138 and anoutlet 140 for connection tomachine 148. A firsthydraulic line 152 connects theoutlet 140 tomachine 148, and a secondhydraulic line 150 connectsmachine 148 to theinlet 138 ofmachine 66 via acooler 156. A thirdhydrostatic line 154 is provided to connectlines - In the
transmission 60ahydraulic machine 66 is permanently configured as a pump andmachine 148 as a motor. At all times during engine operation thepump 66 delivers pressurised working fluid to themotor 148 to drive the aircraft hydraulic systems. - The addition of
transmission 60a to the engine of Figure 2 enables the aircrafthydraulic pump 160 to be positioned closer to the hydraulic systems it drives. By moving thepump 160 from theaccessory gearbox 44 to thepylon structure 52 considerable pipework and hence weight can be removed from the engine. - Referring now to Figure 4 which shows the same gas turbine engine as Figure 2, but with a
transmission system 60 in accordance with a second embodiment of the invention. - The transmission shown connects only two of the engine spools, the high pressure spool and pressure spool 20 and 16 respectively. As shown the transmission comprises first and
second reduction gearboxes 162 and 164 drivingly connected to the engine's high and low pressure spools.Reduction gearbox 162 is connected to the engine's high pressure spool through step-aside gearbox 42, and reduction gearbox 164 to the engine's low pressure spool at the downstream end ofshaft 24 as in the previous embodiment. Gearbox 164 includes bevel gearing (not shown) to transfer the drive fromshaft 24 to a radial output shaft 166. Output shaft 166 is in two parts and extends from the gearbox 164 to a pylon mountedgearbox 168 through a differential 170. This enables the engine's low pressure spool to be used to drive the aircraft hydraulic systems throughpumps International Patent Application 92/01179. In order to integrate the shaft 166 into the transmission the engine's step-aside andaccessory gearboxes constant speed drive 172 together withshafts gearboxes - The axial
differential drive 172 is of the type commonly used in aircraft electrical power generator systems. These drives are configured so that the majority of power passing through them is passed through a differential gear arrangement, whilst a small proportion is used to drive an integral variable speed hydrostatic transmission. The design of the differential drive is such that it allows two different speed inputs to be mechanically summed to produce a single speed output. The axial differential drive is arranged so that one of speed inputs is the mechanical input to the differential gear itself, and the other the output from the hydrostatic transmission. The drive ratio of the differential drive is infinitely variable by virtue of the variable hydrostatic transmission output. - Accordingly the
transmission 60 of the second preferred embodiment operates as follows. During normal engine operation no torque is transferred between engine shafts, the ratio ofdrive 172 essentially follows the ratio of low pressure spool to high pressure spool speed. The only time thedrive 172 deviates from this is following an in-flight combustion flame out condition. Once detected by the engine control (not shown) the ratio of thedrive 172 will alter in accordance with power transfer requirements. As previously discussed this will enable the engine's low pressure spool to accelerate the high pressure spool to the re-light condition. - It will be appreciated, of course, that the
transmission system 60 could be modified further so that power could be transferred selectively between any two of the engine spools. This could be achieved for example by adding a further axial differential drive gear between the engines high pressure 20 and intermediate pressure spool 18, or alternatively a hydrostatic transmission as in the first preferred embodiment. - Turning now to Figure 5 which schematically represents the gas turbine engine of Figure 1 having a transmission system in accordance with a third embodiment of the invention.
- As shown the engine includes all the components of the prior art transmission. The high pressure spool 20 is adapted to drive the
engine accessory gearbox 44 viadrive shafts gearbox 42 as before, and theradial drive shaft 40 is bevelled to the engine's high pressure spool at 180 in the usual way. - In addition, however, the transmission shown in Figure 5 comprises a
reduction gearbox 182 located inregion 68 and drivingly connected to the engine's intermediate pressure spool 18, reduction and bevel gearing 184 drivingly connected to the engine's low pressure spool 16 at the downstream end ofshaft 24, and bevel gearing 186 positioned radially outwards of gearing 184 inregion 68.Gearbox 182 is connected to the engine's intermediate pressure spool 18 by aradial drive shaft 188 bevelled toshaft 32 at 190, and to step asidegearbox 42 by an axial drive shaft 192. In a like manner gearing 186 is connected to gearing 184 by aradial drive shaft 194, and to step-aside gearbox 42 by aninclined drive shaft 196. A clutch assembly 198 is provided in drive shaft 192 so that the drive between the engine's high pressure spool 20 and intermediate pressure spool 18 may be selectively engaged and disengaged. Likewise a secondclutch assembly 200 is provided indrive shaft 196 so that the drive between the engine's high pressure spool 20 and low pressure spool 16 may be selectively engaged and disengaged. - During normal engine operation the
transmission 60 of Figure 5 will operate in the same manner as the transmission system of the prior art. Withclutch assemblies 198 and 200 disengaged the only drive to accessory gearbox is that taken from the engine's high pressure spool. However, by selective engagement and disengagement ofclutch assemblies 198 and 200 it will be seen that power may be transferred between the engine spools, and that the accessory gearbox may be driven by a spool other than the engine's high pressure spool. - From the foregoing it will be seen that engine re-light performance may be improved by engagement of
clutch assembly 200 and disengagement of clutch assembly 198 following a flame-out condition. It will be seen also that engine part speed performance may be improved by engagement of clutch assembly 198 and disengagement ofclutch assembly 200 during part load operation, and ground starting performance by similar engagement and disengagement prior to engine ignition. - A
power transmission system 60 in accordance with a first aspect of a fourth embodiment of the invention is shown in Figure 6. Like Figure 5, Figure 6 shows schematically the gas turbine engine of Figure 1 having all the components of the prior art transmission. The transmission system shown also includes a number of additional components also found in the third embodiment of the invention. - As shown the transmission system of Figure 6 further comprises a first
electrical machine 202 drivingly connected to the engine'saccessory gearbox 44, a secondelectrical machine 204 drivingly connected togearbox 182 and a thirdelectrical machine 206 drivingly connected to the engine's low pressure spool viagearing 208. Theelectrical machines - In the embodiment shown the three
machines control 210 viaseparate lines control 210 comprises switchgear adapted to isolate each of themachines - In this embodiment power may be selectively transferred from the engine's low pressure spool 16 to it's high pressure spool 20 by operating
machine 206 as a generator andmachine 202 as a motor powered by current from thegenerator 206. Alternatively power may be transferred from the engine's intermediate pressure spool 18 to it's high pressure spool 20 by operatingmachine 204 as a generator andmachine 202 as a motor powered by thegenerator 204. Similarly power may be transferred from the engine's high pressure spool 20 to it's intermediate pressure spool 18 by operatingmachines - A second aspect of this embodiment of the invention is shown in Figure 7. In this aspect of the invention the
electrical induction machines electrical machines machines rotors 228 ofmachines engine shafts - The
machines separate lines - Referring now to Figure 8 which shows a power transmission system according to a fifth embodiment of the invention. Like Figures 5 and 6, Figure 8 shows schematically the gas turbine engine of Figure 1 having all the components of the prior art transmission.
- The transmission system shown also comprises an
auxiliary air turbine 230 drivingly connected to theengine accessory gearbox 44. Theauxiliary turbine 230 includes aninlet 232 which is in fluid flow communication with the downstream end ofcompressor 28. Ducting 234 links theinlet 232 of the turbine to an engine bleed flow means 236 positioned at the outlet tocompressor 28. A bleedflow control valve 238 is positionedadjacent compressor bleed 236 so that compressor discharge air may be bled selectively from the engine to drive theturbine 232. - During normal engine operation bleed
flow control valve 238 remains closed and the engine functions as normal. At part speed operation, however, bleedflow control valve 238 opens and compressor discharge air is fed to theturbine 230. The increase in effective discharge area causes the load on theturbine section 30 to increase the thereby the rotational speed of the intermediate pressure spool 18 to fall. This contrasts with an increase in the high pressure spool rotational speed due to the work input fromturbine 230 via the engine accessory drive. - Accordingly the transmission system shown in this embodiment enables power to be selectively transferred from the intermediate pressure spool to high pressure spool by selective operation of
valve 238. Although not shown, the same effect could be achieved by adapting theturbine 232 to be driven by bleed gases bled from the inlet toturbine 30 instead of the outlet tocompressor 28. In such an embodiment opening the bleed toturbine 232 would effectively reduce the area ratio of theturbine 30 and thereby reduce the rotational speed of the spool 18 whilst at the same time increasing the rotational speed of spool 20. - The arrangement shown in Figure 5 offers a number of advantages over such alternative arrangements. In particular it enables the bleed flow generated as a result of flow control at part speed operation to be utilised in a more useful manner. In conventional arrangements the bleed flow is normally dumped to the bypass flow and as such has only a minor effect on cycle efficiency. The arrangement described however enables the bleed flow to be utilised to increase the flow capacity of the engines
high pressure compressor 34 and thereby improve part speed engine performance. - Whilst described with reference to a three shaft gas turbine engine it will be appreciated that all embodiments of the invention are equally applicable to any multispool engine. In particular the invention may be utilised to improve the characteristics of a two shaft engine in much the same way as for the three shaft engine described. Moreover the part speed performance benefits would also be realised in a ground based gas turbine engine incorporating the present invention as well as for aircraft mounted gas turbine engines.
Claims (31)
- A multispool gas turbine engine of the kind comprising at least two independently rotatable engine spools (18,20) and a plurality of power transfer machines (62,64) coupled to the engine spools (18,20), including a first power transfer machine (62) coupled to a first engine spool (20), said machine (62) being capable of operation as a generator (forward power transfer) to take power from the spool (20) or as a motor (reverse power transfer) to drive the spool (20), characterised in that there is provided at least one further power transfer machine (64) coupled to another of the engine spools (18), which further machine (64) is also capable of operation either as a generator (forward power transfer) to take power from the spool (18) or as a motor (reverse power transfer) to drive the spool (18), power transmission means (6) arranged to interconnect said plurality of power transfer machines (62,64), and means (86) to control the operation of said machines (62,64) to transfer power selectively from at least one (20) of the engine spools to at least one other (18) of the engine spools.
- A gas turbine engine according to claim 1 further characterised that the engine (10) includes a plurality of engine driven accessories (44,42) and the transmission means (60) is adapted to transmit power from at least one (20) of the engine spools to drive the accessories.
- A gas turbine engine according to claims 1 or 2 for mounting on an aircraft having a plurality of engine driven accessories (44,42) and wherein the transmission means (60) is further adapted to transmit power from at least one (20) of the engine spools to drive the aircraft accessories.
- A gas turbine engine according to claims 1 to 3 wherein the power transmission means comprises at least two hydraulically coupled flow displacement machines (62,64,66) each drivingly connected to a different one of the engine spools (16,18,20), and whereby at least one of the flow displacement machines is adapted to operate as a motor powered by working fluid energised by at least one other flow displacement machines operating as a pump.
- A gas turbine engine according to claim 4 wherein the flow displacement machines (62,64,66) are variable flow machines.
- A gas turbine engine according to claim 5 wherein the flow displacement machines (62,64,66) are adapted to operate in a first mode as a pump and in a second mode as a motor.
- A gas turbine engine according to claim 6 wherein the transmission means (60) further comprises a control valve means (144,146) for controlling the flow of working fluid between flow displacement machines (62,64,66).
- A gas turbine engine according to claim 6 wherein each of the flow displacement (62,64,660 machines is connected to a respective one of the engine spools (16,18,20) through a gearbox (42,74,78).
- A gas turbine engine according to any one of claims 4 to 8 wherein the working fluid comprises engine oil diverted from the engine oil system.
- A gas turbine engine according to claim 4 wherein a one of the flow displacement machines (66) is connected to an engine low pressure spool (16) and another of the flow displacement machines (62) is connected to an engine high pressure spool (20).
- A gas turbine engine according to claim 10 wherein the low pressure engine spool (16) comprises a fan (12) and a low pressure turbine (22) interconnected by means of a load transmitting shaft (24) and the flow displacement machine (66) is drivingly connected to the downstream end of the load transmitting shaft.
- A gas turbine engine according to any one of claims 4 to 11 wherein a flow displacement machine (62,64,66) is hydraulically coupled to an engine to aircraft accessory power transfer means (40,42,44).
- A gas turbine engine according to claim 10 wherein a further flow displacement unit (64) is connected to an engine intermediate pressure spool (18).
- A gas turbine engine according to claims 1 to 3 wherein the power transmission means (60) comprises at least one variable ratio differential drive means (172) adapted and arranged to transfer power between two of the engine spools (16,18,20).
- A gas turbine engine according to claim 14 wherein the engine (10) comprises a low pressure spool (16) and a high pressure spool (20) and the transmission means (60) comprises at least one variable ratio differential drive means (172) connecting the low pressure spool (16) to the high pressure spool (20).
- A gas turbine engine according to claim 15 wherein the low pressure spool (16) comprises a fan (12) and a low pressure turbine (22) interconnected by means of a load transmitting shaft (24) and the differential drive means (172) is drivingly connected to the low pressure spool (16) through a gearbox (170) connected to the downstream end of the load transmitting shaft (24).
- A gas turbine engine according to claim 16 wherein the gearbox (170) is additionally drivingly connected to an aircraft accessory gearbox (168).
- A gas turbine engine according to claims 1 to 3 wherein the power transmission means (60) comprises a first gearbox means (182) drivingly connected to at least two of the engine spools (18,20).
- A gas turbine engine according to claim 18 wherein the transmission means (60) includes clutch means (198) positioned between at least one of the spools (18) and the gearbox means (182) thereby to provide for selective power transfer between the engine spools (18,20).
- A gas turbine engine according to claim 19 wherein the engine (10) comprises a low pressure spool (16) and a high pressure spool (20) and the first gearbox means (182) is drivingly connected to an engine accessory gearbox means (42,44), and whereby the high pressure spool (20) is adapted to drive the accessory gearbox means (42,44) through the first gearbox means (182) independently of clutch means engagement (198).
- A gas turbine engine according to claims 1 to 3 wherein the power transmission means (60) comprises at least one electrical machine (206) drivingly connected to an engine spool (16) and adapted to operate as a generator, and at least one electrical machine (204) drivingly connected to a further one of the engine spools (18) and adapted to operate as a motor powered by current from the generator (206).
- A gas turbine engine according to claim 21 wherein the electrical machines (204,206) are electrical induction machines each adapted to operate in a forward mode as a motor and in a reverse mode as a generator.
- A gas turbine engine according to claim 22 wherein each of the electrical induction machines (204,206) are drivingly connected to their respective engine spools through reduction gearing (182,208).
- A gas turbine engine according to claim 21 wherein the electrical machines (204,206) are switched reluctance machines each adapted to operate in a forward mode as a motor and in a reverse mode as a generator.
- A gas turbine engine according to claim 22 wherein the rotors of the switched reluctance machines are integral with the respective engine spools.
- A gas turbine engine according to claims 1 to 3 wherein the power transmission means comprises an engine bleed flow means (236) and an auxiliary turbine (230) drivingly connected to at least one of the engine spools (34,36,38), the auxiliary turbine (230) being adapted to receive engine flow from the bleed flow means (236) such that, in use, engine bleed flow may be used to drive said turbine (230) and thereby augment the power of the engine spool.
- A gas turbine engine according to claim 26 wherein the bleed flow means (236) is adapted to deliver air from the compressor section (28) of at least one of the engine spools to the auxiliary turbine (230).
- A gas turbine engine according to claim 27 wherein the auxiliary turbine (230) receives air from one engine spool (28,30,32) and drivingly connects to another engine spool (34,36,38).
- A gas turbine engine according to claims 26 to 28 wherein the engine (10) comprises a low pressure spool (12,22,24) and a high pressure spool (34,36,38) and the auxiliary turbine (230) is drivingly connected to the engine high pressure spool.
- A gas turbine engine according to claims 26 to 29 wherein the auxiliary turbine (230) is drivingly connected to said spool (34,36,38) through reduction gearing (42,44).
- A gas turbine engine according to any preceding claim wherein the engine (10) is a ducted fan engine.
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
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GB9313905 | 1993-07-06 | ||
GB939313905A GB9313905D0 (en) | 1993-07-06 | 1993-07-06 | Shaft power transfer in gas turbine engines |
PCT/GB1994/001450 WO1995002120A1 (en) | 1993-07-06 | 1994-07-05 | Shaft power transfer in gas turbine engines |
Publications (2)
Publication Number | Publication Date |
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EP0659234A1 EP0659234A1 (en) | 1995-06-28 |
EP0659234B1 true EP0659234B1 (en) | 1997-12-29 |
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Application Number | Title | Priority Date | Filing Date |
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EP94919761A Expired - Lifetime EP0659234B1 (en) | 1993-07-06 | 1994-07-05 | Shaft power transfer in gas turbine engines |
Country Status (7)
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US (1) | US5694765A (en) |
EP (1) | EP0659234B1 (en) |
JP (1) | JPH08501370A (en) |
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GB (1) | GB9313905D0 (en) |
RU (1) | RU95112040A (en) |
WO (1) | WO1995002120A1 (en) |
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DE102020117254A1 (en) | 2020-06-30 | 2021-12-30 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine engine and aircraft with a gas turbine engine |
DE102020117255A1 (en) | 2020-06-30 | 2021-12-30 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine engine and aircraft with a gas turbine engine |
Also Published As
Publication number | Publication date |
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US5694765A (en) | 1997-12-09 |
RU95112040A (en) | 1996-12-27 |
DE69407555D1 (en) | 1998-02-05 |
EP0659234A1 (en) | 1995-06-28 |
JPH08501370A (en) | 1996-02-13 |
WO1995002120A1 (en) | 1995-01-19 |
GB9313905D0 (en) | 1993-08-25 |
DE69407555T2 (en) | 1998-08-13 |
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